Roughened coatings for gas turbine engine components

ABSTRACT

A gas turbine engine component with an aluminide coating on at least a portion of an airflow surface that includes a roughening agent effective to provide a desired surface roughness and a deposition process for forming such aluminide coatings. A layer including a binder and the roughening agent may be applied to the superalloy substrate of the component and the aluminide coating formed on the airflow surface portion by exposing the component and layer to an appropriate deposition environment. Suitable roughening agents include metal and ceramic particles that are dispersed on the airflow surface portion before exposure to the deposition environment. The particles, which are substantially intact after the aluminide coating is formed, are dispersed in an effective number to supply the desired surface roughness.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a divisional of U.S. application Ser. No.12/093,980, filed May 16, 2008, which is the National Stage ofInternational Application No. PCT/US2006/006644, filed Feb. 24, 2006.The content of each of these applications is hereby incorporated byreference herein in its entirety for all purposes.

FIELD OF THE INVENTION

The present invention relates to coated metal components and, moreparticularly, gas turbine engine components with a roughened coating andmethods of forming such roughened coatings on gas turbine enginecomponents.

BACKGROUND OF THE INVENTION

Intermetallic layers and coatings are often formed on a surface of ametal component to protect the underlying metal substrate of thecomponent and to extend its useful life during operation. For example,many superalloy components in gas turbine engines, like turbine blades,vanes, and nozzle guides, include an aluminide coating on airflowsurfaces that protects the underlying superalloy base metal from hightemperature oxidation and corrosion. Among other applications, gasturbine engines are used as aircraft or jet engines, such as turbofans.Gas turbine engines are also used in electromotive power generationequipment, such as industrial gas turbine engines, to generateelectricity, and as power plants providing motive forces to propelvehicles.

Generally, gas turbine engines include a compressor for compressing air,a combustor for mixing the compressed air with fuel, such as jet fuel ornatural gas, and igniting the mixture, and a turbine blade assembly forproducing power. In particular, gas turbine engines operate by drawingair into the front of the engine. The air is then compressed, mixed withfuel, and combusted. Hot exhaust gases from the combusted mixture passthrough a turbine, which causes the turbine to spin about an axialcenter and thereby powers the compressor. Aircraft gas turbine engines,referred to herein as jet engines, propel the attached aircraft inresponse to the thrust provided by the flow of the hot exhaust gasesfrom the gas turbine engine. Rotation of the turbine in industrial gasturbine engines generates electrical power and motive power forvehicles.

Gas turbine engines include turbine blades shaped as airfoils andcoupled to the turbine. The hot exhaust gases from the combustor flowover and under each turbine blade. Because of the airfoil shape, theflow path across the top of the airfoil or convex side is much longerthan the flow path underneath the concave side of the turbine blade. Theresult is an aerodynamic lift, which drives each of the turbine bladesin the desired direction. Work is then extracted from the coordinatedrotation of the turbine blades about the axial center of the gasturbine.

Conventional approaches for optimizing aerodynamic lift generated by thespinning turbine blades rely on increasingly radical airfoil shapes andthree-dimensional topologies. However, these conventional approachesthat focus solely upon advances in component geometry introducecomplexity into the component manufacture process and are ultimatelylimited in the improvement in aerodynamic efficiency.

Accordingly, there is a need for gas turbine engine components withimproved lift and methods of forming such gas turbine engine componentsthat avoids the necessity of a complex airfoil shape.

SUMMARY OF INVENTION

The present invention provides, in one aspect, an airflow surface of agas turbine engine component is at least partially covered with analuminide coating including an effective number of substantially-intactparticles dispersed therein such that the aluminide coating has adesired or targeted surface roughness. The gas turbine engine componentis formed from a superalloy material, such as a nickel-based superalloy.The gas turbine engine component may be a turbine blade for a gasturbine engine and, in particular, a jet engine turbine blade for a jetturbine engine.

Advantageously, the aluminide coating on the airflow surface portion maybe formed by a deposition process that includes dispersing the particleson at least the portion of the airflow surface and then forming thealuminide coating that includes the dispersed particles in asubstantially intact condition and in an effective number such that thealuminide coating has a desired or targeted surface roughness. Themethod may include applying a layer containing silicon and theparticles, such as a mixture of silane and either ceramic or metallicparticles, to at least the portion of the airflow surface. Afterapplying the layer, the gas turbine engine component is exposed to adeposition environment effective to form the aluminide coating with thedispersed particles. One suitable deposition environment relies onvaporizing a donor material including a metal effective to form thealuminide layer, which includes the metal from the donor material,silicon from the layer, and the particles from the layer.

The surface finish of the present invention deviates from conventionalturbine blade designs that want the surface finishes on the entireairflow surface to be substantially identical. In contrast, the presentinvention provides a surface finish on one portion of the airflowsurface (i.e., the convex airflow surface found on most gas turbineblades) that differs from the surface finish on another portion of theair flow surface (e.g., the opposite concave airflow surface found onmost gas turbine blades).

The surface finish of the present invention deviates from conventionalturbine blade designs that specify the turbine blades to be as smooth aspossible to contribute to laminar flow and to optimize the flow of hotexhaust gases beneath the concave portion of the airflow surface of theturbine blade. Typically, a desired surface roughness (R_(A)) for thesurface finishes of the convex and concave portions of the airflowsurface is less than about 68 microinches, after aluminiding. Incontrast, the present invention advantageously applies an aluminidecoating to the convex airflow surface portion that increases the surfaceroughness above this conventional desired value. The concave airflowsurface portion may have a conventional surface roughness but, in anyevent, has a smoother surface than all or part of the convex airflowsurface portion. The difference in surface roughness slows the airflowvelocity across the convex airflow surface portion in comparison to theairflow velocity across the concave airflow surface portion.

The present invention improves the aerodynamic efficiency of gas turbineengine components providing aerodynamic lift without the need forcomplex component geometries and/or improves the aerodynamic lift incomponents having complex geometries beyond the gains provided solely bythe geometry.

These and other benefits and advantages of the present invention shallbe made apparent from the accompanying drawings and description thereof.

BRIEF DESCRIPTION OF THE DRAWINGS

The accompanying drawings, which are incorporated in and constitute apart of this specification, illustrate an embodiment of the inventionand, together with a general description of the invention given above,and the detailed description of the embodiment given below, serve toexplain the principles of the invention.

FIG. 1 is a perspective view of a gas turbine engine component with aliquid being applied to a portion of the gas turbine engine component inaccordance with the principles of the present invention;

FIG. 1A is a diagrammatic cross-sectional view of a portion of thecoated gas turbine engine component of FIG. 1;

FIG. 2 is a schematic view showing gas turbine engine components, suchas that from FIG. 1, in a deposition environment of a simple CVDdeposition system for purposes of explaining the principles of thepresent invention;

FIGS. 3A-C are perspective views similar to FIG. 1 in accordance withalternative embodiments of the invention;

FIG. 4 is a diagrammatic cross-sectional view of a portion of a coatedgas turbine engine component of the present invention; and

FIG. 4A is a diagrammatic cross-sectional view of a portion of a coatedgas turbine engine component in accordance with an alternativeembodiment of the present invention.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

With reference to FIG. 1 and in accordance with the principles of thepresent invention, a gas turbine engine component 10, in arepresentative construction, includes an airfoil segment 12 designed tobe in the high-pressure, hot airflow path (as indicated by arrows 13).Integral with airfoil segment 12 is a root 14 used to secure gas turbineengine component 10 to the turbine disk (not shown) of the gas turbineengine (not shown). The airfoil segment 12 is fabricated from anysuitable nickel-, cobalt-, or iron-based high temperature superalloyfrom which such gas turbine engine components 10 are commonly made. Thebase element, typically nickel or cobalt, is by weight the singlegreatest element in the superalloy. For example, where the component 10is a gas turbine component in a jet engine, segment 12 may be thenickel-based superalloy Inconel 795 Mod5A or CMSX-4. The presentinvention is, however, not intended to be limited to any particular gasturbine engine component 10, and may be any high pressure turbine blade,low pressure turbine blade, or any other component of a gas turbinehaving an airfoil surface that generates lift while operating in a jetengine or while operating in an industrial gas turbine engine.

A surface 16 of the airfoil segment 12 of gas turbine engine component10 is divided into airflow surfaces 18, 20 extending between a curvedtip edge 22 and a curved foil tip edge 24. Cooling channels or passagesinternal to airfoil segment 12 include surface cooling holes 26 onsurface 16 so as to permit cooling air to pass through the interior ofairfoil segment 12 while gas turbine engine component 10 is in serviceon the gas turbine engine. The root 14 includes a contoured surface 28extending beneath a platform 30 and is separated from the airfoilsegment 12 by the platform 30.

Depending upon the use of the gas turbine engine component 10,combustion gases in the airflow path 13 may have a temperature as highas 3000° F. This promotes heating of the airfoil segment 12. Gas coolingof the airfoil segment 12 limits operating temperatures to 1800° F. orless. When the gas turbine engine component 10 is in service, portionsof the component 10 below the platform 30 are cooler than the airfoilsegment 12 and, frequently, are at an operating temperature of less than1500° F. when the component 10 is in service. The cooler portionsinclude the root 14, which is coupled with an air-cooled turbine disk ofthe gas turbine.

Airflow surface 20 has a concave shape extending between leading edge 22and trailing edge 24 and airflow surface 18 has a convex shape extendingbetween edges 22 and 24. When the gas turbine engine component 10 iscoupled with a gas turbine engine (not shown) and rotated, leading edge22 is the first to encounter the hot exhaust gases and air in theairflow path 13. The airflow path 13 will split at edge 22. A portion ofthe hot exhaust gases and air in air flow path 13 will flow acrossairflow surface 18 and another portion of the hot exhaust gases and airwill flow across airflow surface 20. Due to the difference in curvatureand length, the flow velocity will be greater across airflow surface 18than across airflow surface 20. Due to the familiar Bernoulli'sprinciple, lift is generated because the pressure is greater nearairflow surface 20 than near airflow surface 18. The split airflowrecombines after passing trailing edge 24.

With reference to FIGS. 1 and 1A, a layer 40 is applied to all or aportion of convex airflow surface 18 of gas turbine engine component 10before an aluminide coating 42 (FIG. 4) is formed on selected regions ofconvex airflow surface 18 in a CVD apparatus 50 (FIG. 2). The layer 40may be applied as a liquid or solution and then dried to remove solventand form a solid or semi-solid coating on the gas turbine enginecomponent 10 before aluminiding occurs.

The layer 40 is applied to all or a portion of convex airflow surface18, such as by hand application with a paintbrush B or another type ofconventional applicator recognized by a person having ordinary skill inthe art. Alternatively, gas turbine engine component 10 may be sprayedwith a suitable liquid or solution before drying and aluminiding.Thereafter, the coated gas turbine engine component 10 (which mayadvantageously first be dried and heated) is placed into a depositionenvironment 52 (FIG. 2) whereupon the aluminide coating 42 will beformed on convex airflow surface 18 to the desired thickness.

The layer 40 applied to all or a portion of the convex airflow surface18 is initially a liquid or solution that includes a binder 38 and aroughening agent, such as inorganic particles 44, blended with thebinder 38. The liquid forming the binder 38 may be a silicon-containingbinder such as a silane and, advantageously, may be a high-viscositysilane. Silanes suitable for use in the present invention may havemono-, bis-, or tri-functional trialkoxy silane. The silane may be abifunctional trialkoxy silyl, preferably trimethoxy, or triethoxy silylgroups. Amino silanes may also be used, although thio silanes may not bedesired due to their sulfur content. Bisfunctional silane compounds arewell known, and two suitable for use in the present invention arebis(triethoxysilyl) ethane and bis(trimethoxysilyl) methane. In both ofthese compounds, the bridging group between the two silane moieties isan alkyl group. Additional commercially available silanes include, butare not limited to,

-   -   1,2-Bis(tetramethyldisoloxanyl) Ethane    -   1,9-Bis(triethoxysilyl) Nonane    -   Bis(triethoxysilyl) Octane    -   Bis(trimethoxysilyl Ethane    -   1,3-Bis(trimethylsiloxy)-1,3-Dimethyl Disiloxane    -   Bis(trimethylsiloxy) Ethylsilane    -   Bis(trimethylsiloxy) Methylsilane    -   A1-501 available from AG Chemetall (Frankfurt Germany)

The silane of binder 38 may be neat, in an aqueous solution, or in anaqueous/alcohol solvent solution. A solvent for the latter type ofsolution may contain from about 1% to 2% by volume (vol. %) to about 30vol. % deionized water with the remainder of the solution being a loweralcohol, such as methanol, ethanol, isopropanol, or the like. Thesolvent is combined with the silane and glacial acetic acid to establisha pH of about 4 to 6. The concentration of the silane compound is notrelevant as long as the silane remains in solution during application.Generally, the solution will include about 1% to about 20% silane, whichmay be measured either by volume or by weight in this concentrationrange.

The binder 38 of layer 40 applied to gas turbine engine component 10 isallowed to dry and then is heated, such as with a heat gun (not shown)or in a heated enclosure (not shown), to a temperature suitable torelease or remove solvent from the binder 38 and provide a solid orsemisolid cured state. Before curing, the layer 40 on the convex airflowsurface 18 may first be allowed to dry, such as underneath a lamp (notshown), to partially remove the constituent solvent. Generally, thelayer 40 is applied in an amount of about 0.01 g/cm² to about 2.0 g/cm².Multiple layers 40 of liquid or solution may be applied to convexairflow surface 18, each individual layer 40 being dried and heatedbefore applying the next successive layer 40. As used herein, the layer40 may refer to either the initially applied layer of liquid or solutionor, without limitation, to the cured or dried layer that has had solventremoved from binder 38 by heating and/or air curing at room temperature.

The particles 44 of layer 40 constituting the roughening agent mayadvantageously be composed of a ceramic, such as silica, alumina,chromium dioxide, yttria, hafnia, zirconia, and combinations andmixtures thereof. For example, the particles 44 may be a fine aluminaflour having a mesh size on the order of 270 to 325 mesh or finer.Alternatively, the particles 44 may include a metal, such as boron,aluminum, chromium, yttrium, hafnium, zirconium, and combinations andalloys thereof. Alternatively, the particles 44 may be a metallic powdercomprised of metallurgy identical to the base metal constituting asubstrate 46 of the gas turbine engine component 10 and with an optionaladdition of less than about 1% by weight of boron powder. Preferably,the layer 40 is not allowed to infiltrate into the cooling holes 26during application to the gas turbine engine component 10. The binder 38of layer 40, after curing, secures the particles 44 to the airflowsurface 18 during the aluminiding process. The invention contemplatesother types or compositions of binders 38, which may lack a siliconcontent, may be used to retain the dispersed particles 44 on airflowsurface 18 before aluminiding.

With reference to FIG. 2, a CVD apparatus 50 suitable for use in formingthe aluminide coating 42 (FIG. 4) includes a main reaction chamber 54enclosing the interior space defining a deposition environment 52 whenpurged of atmospheric gases, and evacuated. Inert gas, such as argon, issupplied from a gas supply 56 to the reaction chamber 54 through aninlet port 58 defined in the wall of chamber 54. An exhaust port 60defined in the wall of the reaction chamber 54 is coupled with a vacuumpump 62 capable of evacuating the reaction chamber 54 to a vacuumpressure. One or more gas turbine engine components 10 are introducedinto the reaction chamber 54 and are situated away from a source ofextrinsic metal, as explained below.

Positioned within the reaction chamber 54 is a mass or charge of a soliddonor material 64, a mass or charge of an activator material 66, andseveral gas turbine engine components 10. Suitable solid donor materials64 include alloys of chromium and aluminum, which are preferably low insulfur content (<3 ppm sulfur). One suitable donor material 64 is 44 wt% aluminum and balance chromium. Appropriate activator materials 66suitable for use in the invention include, but are not limited to,aluminum fluoride, aluminum chloride, ammonium fluoride, ammoniumchloride, and ammonium bifluoride. The reaction chamber 54 is heated toa temperature effective to cause vaporization of the activator material66, which is transported as diagrammatically indicated by arrows 65within the deposition environment 52 to the solid donor material 64.Typically, this temperature ranges from about 1950° F. to about 2000° F.Interaction between the vaporized activator material 66 and the soliddonor material 64 promotes the release of a vapor phase reactant fromthe solid donor material 64. This vapor contains an extrinsic metal,typically aluminum, that contributes a first extrinsic metal forincorporation into an aluminide coating 42 (FIG. 4) formed on component10, as diagrammatically indicated by arrows 68. The extrinsic metal isseparate, distinct, and independent from the material comprising the gasturbine engine component 10 and any coating preapplied to component 10.

With reference to FIGS. 3A-C in which like reference numerals refer tolike features in FIG. 2 and in accordance with alternative embodimentsof the invention, layer 40 may be provided on only selected regions ofthe convex airflow surface 18. As shown in FIG. 3A, layer 40 may beapplied in discrete areas distributed across the convex airflow surface18 as a plurality of substantially-parallel stripes 70 having eithersmooth edges or jagged, uneven edges. The stripes 70 of layer 40 arealigned substantially parallel to the leading and trailing edges 22, 24that bound the airflow surface 18.

As shown in FIG. 3B, layer 40 may be applied across the convex airflowsurface 18 in discrete areas defined by a plurality of discrete islandsor areas 72 arranged either randomly or in specific rows and/or columns.The peripheral boundary surrounding the discrete areas 72 of layer 40may be irregular, as shown, angular, curvilinear, regular (e.g.,circular), or a distribution of different shapes.

As shown in FIG. 3C, the layer 40 may be applied across the convexairflow surface 18 as a pattern of substantially-parallel stripes 74inclined diagonally across the convex airflow surface 18. The resultantairflow path is believed to occur in channels defined between adjacentstripes 74 and is directed toward the root 14 of the component 10, whichrepresents a non-airflow surface. Each of the stripes 74 of layer 40intersects at least one of the first and second edges 22, 24 that boundthe airflow surface 18. This particular pattern for layer 40 may causethe air in airflow path 13 to twist as it tumbles through the gasturbine engine (not shown).

With reference to FIG. 4, the aluminide coating 42 is formed on ametallic substrate 46 of the gas turbine engine component 10 across atleast the airflow surfaces 18, 20. The aluminide coating 42 on theconvex airflow surface 18 on areas with the pre-applied layer 40 willinclude the particles 44 and may include one or more elements from thebinder 38 (FIG. 1A). The spatial distribution of the particles 44determines the topography of the aluminide coating 42 in these areas.Advantageously, the viscosity of the binder 38 is sufficient to coverthe particles 44 so that the particles 44 are buried or submerged in thebinder 38 before aluminiding.

The particles 44 operate to effectively increase the surface roughnessof the aluminide coating 42 in comparison with adjacent portions ofconvex airflow surface 18, if any, lacking layer 40 before aluminiding.Particles 44 create raised or elevated surface irregularities or moundsin the aluminide coating 42 at distributed locations across the convexairflow surface 18. This difference in surface finish is best apparentfrom FIG. 4 as areas of the aluminide coating 42 proximate or local toeach particle 44 will have an average thickness of h₁, as compared tothe nominal thickness, h₀, of the aluminide coating 42 in regionsbetween adjacent particles 44 and not affected by the presence of theparticles 44. Of course, the increase in average surface roughness willalso reflect the number or density of particles 44 and willstatistically include portions of aluminide coating 42 overlying theparticles 44 and having local effective thicknesses ranging between h₁and h₀ due to the mounding. The thickness h₀ may be thicker than thethickness of the aluminide coating 42 formed on other surfaces of thegas turbine engine component 10, such as on concave airflow surface 20,because of the presence of silicon originating from binder 38.

The particles 44 remain substantially intact after the aluminidingprocess forming the aluminide coating 42. Preferably, the particles 44originally dispersed in the pre-applied layer 40 are incorporated intothe aluminide coating 42 without significant degradation by thealuminiding process or at the temperature of the aluminiding process.The number of particles 44 dispersed in the aluminide coating 42 iseffective to provide the aluminide coating 42 with a desired surfaceroughness. The value of the average or peak surface roughness iscontingent upon, among other parameters, the size, shape, distribution,and number of particles 44 dispersed in the aluminide coating 42.Preferably, the surface finish of aluminide coating 42 has an averagesurface roughness (R_(A)) greater than a conventional surface finish,considered to lack particles similar to particles 44, of about 68microinches. Advantageously, the average surface roughness of aluminidecoating 42 is greater than about 75 microinches. More advantageously,the average surface roughness of aluminide coating 42 is greater thanabout 100 microinches. Most advantageously, the particles 44 influencethe aluminide coating 42 to provide an average surface roughness thatranges from about 120 microinches to about 130 microinches.

The particles 44 are illustrated in FIG. 4 as having a substantiallyuniform size. However, the invention is not so limited as particles 44may have a distribution of sizes with a size range effective provide thedesired surface finish on airflow surface 20. The particles 44 areillustrated in FIG. 4 as being approximately spherical. However, theinvention is not so limited as particles 44 may have other appropriatethree-dimensional geometrical shapes, such as elongated cylinders, rods,needles, pyramids, etc. The particles 44 are illustrated in FIG. 4 asbeing positioned approximately at the position of the original concaveairflow surface 18. However, the invention is not so limited asparticles 44 may be positioned with a distribution of locations acrossthe thickness of aluminide coating 42 between the surfaces 18 and 45.

In this specific embodiment of the present invention, aluminide coating42 operates as an environmental coating having a working surface 45exposed to the atmosphere with the gas turbine engine component 10 inservice. The general composition of aluminide coating 42 in regions ofthe convex airflow surface 18 initially covered by layer 40 mayadvantageously include a concentration of silicon if the binder 38contains silicon. In this instance, the concentration of silicon in thealuminide coating 42 may be, for example, about 0.5 percent by weight(wt %).

The presence of silicon in the aluminide coating 42 may also increasethe thickness of the aluminide coating 42 in regions of the convexairflow surface 18 initially covered by layer 40, in comparison with thealuminide coating 42 on regions of the convex airflow surface 18 notinitially covered by layer 40. This increased comparative thickness mayalso effectively contribute to the roughening of the convex airflowsurface 18 if the layer 40 is applied to selected regions, as shown forexample in FIGS. 3A-C.

With reference to FIG. 4A, the invention contemplates aluminide coating42 may partially diffuse into the substrate 46 beneath the originalconvex airflow surface 18 of the substrate 46, instead of being a purelyadditive layer as shown in FIG. 4. The resulting aluminide coating 42includes a diffusion region 41 that extends beneath the formed positionof the original convex airflow surface 18 and an additive region 43overlying the former position of the original convex airflow surface 18.In this instance, the outermost boundary of the additive region 43defines the working surface 45 of aluminide coating 42 when the gasturbine engine component 10 is in service. Additive region 43 is analloy that includes a relatively high concentration of the donor metalaluminum and a concentration of a metal, for example nickel, fromsubstrate 46 outwardly diffusing from component 10. By contrast,diffusion region 41 has a lower concentration of aluminum and arelatively high concentration of the metal of substrate 46.

The present invention may be used in combination with the application ofa platinum aluminide coating on gas turbine engine component 10. In thisinstance, layer 40 is placed on the gas turbine engine component 10after the coating of platinum but before aluminiding.

The aluminide layer 42 containing particles 44 may also be formed on gasturbine engine components 10 including the silicon-containing layer 30by various alternative techniques known in the art, including but notlimited to dynamic CVD and pack coating deposition processes such as anabove-the-pack process or an in-the-pack process or by electrosparkdeposition or alloying.

The present invention is generally applicable to turbine enginecomponents 10 used in the gas turbines of jet engines, the gas turbinesof industrial gas turbine engines, or in other turbomachinery. Inparticular, the present invention is applicable for roughening turbineblades in such engines and, more particularly, for roughening turbineblades in the gas turbines used in jet engines.

While the present invention has been illustrated by the description ofan embodiment thereof and specific examples, and while the embodimenthas been described in considerable detail, it is not intended torestrict or in any way limit the scope of the appended claims to suchdetail. Additional advantages and modifications will readily appear tothose skilled in the art. The invention in its broader aspects istherefore not limited to the specific details, representative apparatusand methods and illustrative examples shown and described. Accordingly,departures may be made from such details without departing from thescope or spirit of applicant's general inventive concept.

1. A deposition process for a superalloy gas turbine engine componenthaving an airflow surface, comprising: dispersing a plurality ofparticles on at least a portion of the airflow surface; and forming analuminide coating on the airflow surface portion that includes thedispersed particles in a substantially intact condition and in aneffective number such that the aluminide coating has a desired surfaceroughness.
 2. The deposition process of claim 1 wherein forming thealuminide coating further comprises: exposing the gas turbine enginecomponent to a deposition environment effective to form the aluminidecoating.
 3. The deposition process of claim 2 wherein dispersing theparticles further comprises: applying a layer comprising the particleson the airflow surface portion and a binder effective to adhere theparticles to the airflow surface while the aluminide coating is formed.4. The deposition process of claim 3 wherein exposing the gas turbineengine component to the deposition environment further comprises:forming the aluminide coating from a metal originating from a vaporizeddonor material, silicon from the layer, and the particles from thelayer.
 5. The deposition process of claim 1 wherein dispersing theparticles further comprises: applying a layer comprising the particleson the airflow surface portion and a binder effective to adhere theparticles to the airflow surface while the aluminide coating is formed.6. The deposition process of claim 5 wherein applying the layer furthercomprises: applying the layer across an entire surface area of theairflow surface.
 7. The deposition process of claim 5 wherein applyingthe layer further comprises: applying the layer in a discrete areas onthe airflow surface.
 8. The deposition process of claim 5 wherein thesubstrate includes first and second edges bounding the airflow surface,and applying the layer further comprises: applying stripes of the layerthat are inclined to intersect at least one of the first and secondedges.
 9. The deposition process of claim 5 wherein the substrateincludes first and second edges bounding the airflow surface, andapplying the layer further comprises: applying stripes of the layer thatare substantially parallel to the first and second edges.
 10. Thedeposition process of claim 5 wherein applying the layer furthercomprises: mixing the particles with a binder comprising a silanesolution; and placing the binder on the airflow surface portion so as tobind the particles to the airflow surface portion.
 11. The depositionprocess of claim 1 wherein the desired surface roughness of thealuminide coating is greater than about 0.68 microinches.
 12. Thedeposition process of claim 1 wherein the desired surface roughness ofthe aluminide coating is greater than about 0.75 microinches.
 13. Thedeposition process of claim 1 wherein the desired surface roughness ofthe aluminide coating is greater than about 100 microinches.
 14. Thedeposition process of claim 1 wherein the desired surface roughness ofthe aluminide coating ranges from about 120 microinches to about 130microinches.
 15. The deposition process of claim 1 wherein the airflowsurface has a convex curvature.
 16. The deposition process of claim 1wherein the particles are distributed in separated discrete areas acrossthe airflow surface.